WO2014092834A2 - High pressure rotor disk - Google Patents
High pressure rotor disk Download PDFInfo
- Publication number
- WO2014092834A2 WO2014092834A2 PCT/US2013/061319 US2013061319W WO2014092834A2 WO 2014092834 A2 WO2014092834 A2 WO 2014092834A2 US 2013061319 W US2013061319 W US 2013061319W WO 2014092834 A2 WO2014092834 A2 WO 2014092834A2
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- ratio
- recited
- gas turbine
- bore
- diameter
- Prior art date
Links
- 238000004519 manufacturing process Methods 0.000 claims description 5
- 238000004891 communication Methods 0.000 claims description 4
- 239000012530 fluid Substances 0.000 claims description 4
- 238000000034 method Methods 0.000 claims description 4
- 239000007789 gas Substances 0.000 description 38
- 239000000446 fuel Substances 0.000 description 6
- 239000000463 material Substances 0.000 description 3
- PXHVJJICTQNCMI-UHFFFAOYSA-N Nickel Chemical compound [Ni] PXHVJJICTQNCMI-UHFFFAOYSA-N 0.000 description 2
- 230000004323 axial length Effects 0.000 description 2
- 238000003754 machining Methods 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 229910000975 Carbon steel Inorganic materials 0.000 description 1
- 229910045601 alloy Inorganic materials 0.000 description 1
- 239000000956 alloy Substances 0.000 description 1
- 238000005266 casting Methods 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 238000002485 combustion reaction Methods 0.000 description 1
- 238000012937 correction Methods 0.000 description 1
- 230000007423 decrease Effects 0.000 description 1
- 238000005242 forging Methods 0.000 description 1
- 230000013011 mating Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 229910052759 nickel Inorganic materials 0.000 description 1
- 238000012545 processing Methods 0.000 description 1
- 230000001141 propulsive effect Effects 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
- 238000012546 transfer Methods 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
- F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2200/00—Mathematical features
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/24—Rotors for turbines
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
- Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
- Y10T29/00—Metal working
- Y10T29/49—Method of mechanical manufacture
- Y10T29/49316—Impeller making
- Y10T29/4932—Turbomachine making
- Y10T29/49321—Assembling individual fluid flow interacting members, e.g., blades, vanes, buckets, on rotary support member
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
Abstract
A rotor disk for a gas turbine engine is disclosed and formed to enable operation at high rotational speeds in a high temperature environment. The rotor disk is formed to include a bore, a live rim diameter and an outer diameter related to each other according to defined relationships.
Description
HIGH PRESSURE ROTOR DISK
CROSS REFERENCE TO RELATED APPLICATION
[0001] This application claims priority to U.S. Provisional Application No. 61/707,009 filed on September 28, 2012.
BACKGROUND
[0002] A gas turbine engine typically includes a fan section, and a core engine section including a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The highspeed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
[0003] Turbine and compressor rotor disks operate at high speeds and support blades. Exhaust gases produced in the combustor drive a rotor disk within the turbine section and thereby rotation of a corresponding rotor disk within the compressor section. The turbine disk is attached to drive a shaft that in turn drives the compressor or the fan section.
[0004] Engine manufactures continuously seek improvements to thermal, weight and propulsive efficiencies. Improvements to engine architectures have enabled higher speeds and operation at increased temperatures. Accordingly, it is desirable to develop rotor disks that perform at higher speeds and greater temperatures.
SUMMARY
[0005] A gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a compressor section including a high pressure compressor and a low pressure compressor. A combustor is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. The turbine section includes a high pressure turbine driving the high pressure compressor and a low pressure turbine driving the low pressure compressor. At least one of the high pressure
turbine and the high pressure compressor includes a disk having a bore diameter (D) related to a bore width (W) according to a ratio (D/W) between about 1.25 and about 1.65.
[0006] In a further embodiment of the foregoing gas turbine engine, the ratio (D/W) is between about 1.35 and about 1.55.
[0007] In a further embodiment of any of the foregoing gas turbine engines, the ratio (D/W) is about 1.45.
[0008] In a further embodiment of any of the foregoing gas turbine engines, the ratio (D/W) is equal to 1.45.
[0009] In a further embodiment of any of the foregoing gas turbine engines, the disk includes an outer diameter (OD) related to the bore diameter (D) according to a ratio (OD/D) that is between about 2.95 and about 3.25.
[0010] In a further embodiment of any of the foregoing gas turbine engines, the ratio (OD/D) is between about 3.04 and 3.20.
[0011] In a further embodiment of any of the foregoing gas turbine engines, the ratio (OD/D) is about 3.15.
[0012] In a further embodiment of any of the foregoing gas turbine engines, the ratio (OD/D) is equal to 3.15.
[0013] In a further embodiment of any of the foregoing gas turbine engines, the disk includes a live rim diameter (d) related to the bore diameter (D) according to a ratio (d/D) that is between about 2.25 and about 3.00.
[0014] In a further embodiment of any of the foregoing gas turbine engines, the ratio (d/D) is between about 2.50 and about 2.75.
[0015] In a further embodiment of any of the foregoing gas turbine engines, the ratio (d/D) is about 2.69.
[0016] In a further embodiment of any of the foregoing gas turbine engines, the ratio (d/D) is equal to about 2.69.
[0017] A rotor disk for a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes an outer diameter (OD) related to a bore diameter (D) according to a ratio (OD/D) that is between about 2.95 and about 3.25.
[0018] In a further embodiment of the foregoing rotor disk, the ratio (OD/D) is between about 3.04 and 3.20.
[0019] In a further embodiment of any of the foregoing rotor disks, the ratio (OD/D) is about 3.15.
[0020] In a further embodiment of any of the foregoing rotor disks, the bore diameter (D) is related to a bore width (W) according to a ratio (D/W) between about 1.25 and about 1.65.
[0021] In a further embodiment of any of the foregoing rotor disks, the ratio (D/W) is between about 1.53 and about 1.55.
[0022] In a further embodiment of any of the foregoing rotor disks, the ratio (D/W) is about 1.45.
[0023] In a further embodiment of any of the foregoing rotor disks, the disk includes a live rim diameter (d) related to the bore diameter (D) according to a ratio (d/D) that is between about 2.25 and about 3.00.
[0024] In a further embodiment of any of the foregoing rotor disks, the ratio (d/D) is between about 2.50 and about 2.75.
[0025] In a further embodiment of any of the foregoing rotor disks, the ratio (d/D) is about 2.69.
[0026] A method of fabricating a rotor disk for a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes forming a bore which includes a bore diameter (D) and a live rim diameter (d) with a ratio (d/D) of the live rim diameter (d) to the bore diameter (D) being between about 2.25 and about 3.00, forming at least one lug for mounting a blade at a live rim diameter (d), and forming an outer diameter (OD).
[0027] In a further embodiment of the foregoing method, includes forming the disk to include a ratio (OD/D) of the outer diameter (OD) to the bore diameter (D) between about 2.95 and about 3.25.
[0028] In a further embodiment of any of the foregoing methods, includes forming a bore including a bore diameter (D) and a bore width (W) in a direction parallel to an axis of intended rotation. The bore diameter (D) is related to the bore width (W) according to a ratio (D/W) that is between about 1.25 and 1.65.
[0029] Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
[0030] These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is a schematic view of an example gas turbine engine.
Figure 2 is a cross-section of an example rotor disk for a gas turbine
Figure 3 is a front view of the example rotor disk for a gas turbine engine.
DETAILED DESCRIPTION
[0034] Figure 1 schematically illustrates an example gas turbine engine 20 that includes a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B while the compressor section 24 draws air in along a core flow path C where air is compressed and communicated to a combustor section 26. In the combustor section 26, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24.
[0035] Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three- spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that
enables a high pressure turbine to drive a high pressure compressor of the compressor section.
[0036] The example engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
[0037] The low speed spool 30 generally includes an inner shaft 40 that connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46. The inner shaft 40 drives the fan 42 through a speed change device, such as a geared architecture 48, to drive the fan 42 at a lower speed than the low speed spool 30. The high-speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A.
[0038] A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. In one example, the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54. In another example, the high pressure turbine 54 includes only a single stage. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.
[0039] The example low pressure turbine 46 has a pressure ratio that is greater than about five (5). The pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
[0040] A mid-turbine frame 58 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid- turbine frame 58 further supports bearing systems 38 in the turbine section 28 as well as setting airflow entering the low pressure turbine 46.
[0041] Airflow through the core flow path C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high
pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 58 includes vanes 60, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 58. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
[0042] The disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
[0043] In one disclosed embodiment, the gas turbine engine 20 includes a bypass ratio greater than about ten (10: 1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.
[0044] A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point.
[0045] "Low fan pressure ratio" is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as disclosed herein according to one non- limiting embodiment is less than about 1.50. In another non- limiting embodiment the low fan pressure ratio is less than about 1.45.
[0046] "Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7 °R)]0'5. The "Low corrected fan tip speed", as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.
[0047] The example gas turbine engine includes the fan 42 that comprises in one non-limiting embodiment less than about twenty six (26) fan blades. In another non-limiting embodiment, the fan section 22 includes less than about twenty (20) fan blades. Moreover, in one disclosed embodiment the low pressure turbine 46 includes no more than about six (6) turbine rotors schematically indicated at 34. In another non-limiting example embodiment the low pressure turbine 46 includes about three (3) turbine rotors. A ratio between the number of fan blades 42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of blades 42 in the fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.
[0048] As appreciated, although an engine for mounting and powering an aircraft is described and shown, the present disclosure may also provide benefits to land based and industrial gas turbine engines.
[0049] Referring to Figures 2 and 3, with continued reference to Figure 1, an example rotor disk 62 is shown and includes a plurality of lugs 72 for supporting blades 74 (Figure 3). The example rotor disk 62 is provided as part of the high pressure turbine 54. However, the example rotor disk may also be part of the high pressure compressor 52.
[0050] The rotor disk 62 supports the turbine blades 74 that are driven by high speed exhaust gases generated in the combustor section 26. The example turbine disk 62 includes an outer diameter 64, a live rim diameter 66 and a bore 78 having a bore diameter 70. The bore 78 includes a width 68 in a direction parallel to the axis A.
[0051] The bore diameter 70 is that diameter between an inner most surface of the bore 78 about the axis A. The live rim disk diameter 64 is the diameter that extends between bottom and radially inward surfaces of the disk lugs 72. The turbine blades 74 are supported within the disk lugs 72 by corresponding mating profiles disposed at an interface 76.
[0052] The bore width 68 of the rotor disk 62 in this example is the greatest width on the main body of the rotor disk 62. The greatest width of the main body of the rotor disk 62 does not include additional widths associated with appendages, arms or other structures that extend from the main body of the rotor disk 62. The example bore width 68 is disposed at a distance spaced apart from the axis A determined to provide desired performance properties and to accommodate high rotational speeds encountered during operation. The example distance is defined as the bore diameter 70 (D).
[0053] The speed at which the high pressure turbine rotor disk 62 operates is accommodated at least in part by a relationship between the live rim diameter 66 (d) to the bore diameter 70 (D) defined by a ratio of the live rim diameter 66 to the bore diameter 70 (i.e., d/D). In the disclosed example embodiment the ratio is between about 2.25 and about 3.00. In another disclosed example embodiment, the ratio d/D is between about 2.50 and about 2.75. In another disclosed example embodiment the ratio d/D is about 2.69.
[0054] The disk 62 includes the bore width 68 (W). The bore width 68 (W) is the width at the bore 78 parallel to the axis A. In one non-limiting embodiment, a relationship between the bore width 68 (W) and the bore diameter 70 (D) is defined by a ratio of D/W. In a disclosed example the ratio D/W is between about 1.25 and about 1.65. In another disclosed embodiment the ratio of D/W is between about 1.35 and about 1.55. In a further embodiment the ratio of D/W is about 1.45.
[0055] The outer diameter 64 (OD) is related to the bore diameter 70 (D) by a ratio of the outer diameter 64 (OD) to the bore diameter 70 (D) (i.e., OD/D). In one example the ratio OD/D is between about 2.95 and 3.25. In another embodiment the ratio (OD/D) is between about 3.04 and 3.20. In a further embodiment the ratio (OD/D) is about 3.15.
[0056] The example rotor disk for a gas turbine engine is fabricated by forming a bore including the bore diameter (D) and the bore width (W) in a direction parallel to an axis of intended rotation according to the above disclosed ratio. Additional processing steps are performed to form at least one lug 72 at the live rim diameter 66 (d). Further, the outer diameter 64(OD) is formed according to the above defined ratios. Fabrication further includes forming the rotor disk to include the ratio (OD/D) of the outer diameter (OD) to the bore diameter (D) as disclosed above. The fabrication further includes forming the rotor disk
to include the ratio (d/D) of the live rim diameter (d) to the bore diameter (D) as disclosed above.
[0057] The example rotor disk 62 is fabricated from a material capable of withstanding rotational speeds and temperatures encountered during operation of the gas turbine engine. The rotor disk 62 can be formed from any material or combination of materials such as for example nickel-based alloys and carbon steels. Moreover, it is within the contemplation of this disclosure that the example rotor disk 62 may be fabricated utilizing known fabrication and machining processes verified to enable operation of a completed rotor disk 62 within desired performance parameters within the gas turbine engine. For example, the rotor disk 62 may be fabricated as a casting or forging followed by machining to obtain the desired shape and features.
[0058] The example rotor disk includes relationships that enable performance at high rotational speeds in the high temperature environment of the high pressure turbine 54.
[0059] Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.
Claims
1. A gas turbine engine comprising:
a compressor section including a high pressure compressor and a low pressure compressor;
a combustor in fluid communication with the compressor section;
a turbine section in fluid communication with the combustor, wherein the turbine section includes a high pressure turbine driving the high pressure compressor and a low pressure turbine driving the low pressure compressor, wherein at least one of the high pressure turbine and the high pressure compressor includes a disk having a bore diameter (D) related to a bore width (W) according to a ratio (D/W) between about 1.25 and about 1.65.
2. The gas turbine engine as recited in claim 1 , wherein the ratio (D/W) is between about 1.35 and about 1.55.
3. The gas turbine engine as recited in claim 1, wherein the ratio (D/W) is about 1.45.
4. The gas turbine engine as recited in claim 1, wherein the ratio (D/W) is equal to 1.45.
5. The gas turbine engine as recited in claim 1, wherein the disk includes an outer diameter (OD) related to the bore diameter (D) according to a ratio (OD/D) that is between about 2.95 and about 3.25.
6. The gas turbine engine as recited in claim 5, wherein the ratio (OD/D) is between about 3.04 and 3.20.
7. The gas turbine engine as recited in claim 5, wherein the ratio (OD/D) is about 3.15.
The gas turbine engine as recited in claim 5, wherein the ratio (OD/D) is equal to
9. The gas turbine engine as recited in claim 1, wherein the disk includes a live rim diameter (d) related to the bore diameter (D) according to a ratio (d/D) that is between about 2.25 and about 3.00.
10. The gas turbine engine as recited in claim 9, wherein the ratio (d/D) is between about 2.50 and about 2.75.
11. The gas turbine engine as recited in claim 9, wherein the ratio (d/D) is about 2.69.
The gas turbine engine as recited in claim 9, wherein the ratio (d/D) is equal to about
13. A rotor disk for a gas turbine engine comprising:
an outer diameter (OD) related to a bore diameter (D) according to a ratio (OD/D) that is between about 2.95 and about 3.25.
14. The rotor disk as recited in claim 13, wherein the ratio (OD/D) is between about 3.04 and 3.20.
15. The rotor disk as recited in claim 13, wherein the ratio (OD/D) is about 3.15.
16. The rotor disk as recited in claim 13, wherein the bore diameter (D) is related to a bore width (W) according to a ratio (D/W) between about 1.25 and about 1.65.
17. The rotor disk as recited in claim 13, wherein the ratio (D/W) is between about 1.53 and about 1.55.
18. The rotor disk as recited in claim 13, wherein the ratio (D/W) is about 1.45.
19. The rotor disk as recited in claim 13, wherein the disk includes a live rim diameter (d) related to the bore diameter (D) according to a ratio (d/D) that is between about 2.25 and about 3.00.
20. The rotor disk as recited in claim 19, wherein the ratio (d/D) is between about 2.50 and about 2.75.
21. The rotor disk as recited in claim 19, wherein the ratio (d/D) is about 2.69.
22. A method of fabricating a rotor disk for a gas turbine engine comprising;
forming a bore including a bore diameter (D) and a live rim diameter (d) with a ratio (d/D) of the live rim diameter (d) to the bore diameter (D) being between about 2.25 and about 3.00
forming at least one lug for mounting a blade at a live rim diameter (d); and forming an outer diameter (OD).
23. The method as recited in claim 22, including forming the disk to include a ratio (OD/D) of the outer diameter (OD) to the bore diameter (D) between about 2.95 and about 3.25.
24. The method as recited in claim 22, including forming a bore including a bore diameter (D) and a bore width (W) in a direction parallel to an axis of intended rotation, wherein the bore diameter (D) is related to the bore width (W) according to a ratio (D/W) that is between about 1.25 and 1.65.
Applications Claiming Priority (4)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US201261707009P | 2012-09-28 | 2012-09-28 | |
US61/707,009 | 2012-09-28 | ||
US13/713,257 US10119400B2 (en) | 2012-09-28 | 2012-12-13 | High pressure rotor disk |
US13/713,257 | 2012-12-13 |
Publications (2)
Publication Number | Publication Date |
---|---|
WO2014092834A2 true WO2014092834A2 (en) | 2014-06-19 |
WO2014092834A3 WO2014092834A3 (en) | 2014-08-28 |
Family
ID=50484084
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/US2013/061319 WO2014092834A2 (en) | 2012-09-28 | 2013-09-24 | High pressure rotor disk |
Country Status (2)
Country | Link |
---|---|
US (2) | US10119400B2 (en) |
WO (1) | WO2014092834A2 (en) |
Families Citing this family (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8935913B2 (en) | 2012-01-31 | 2015-01-20 | United Technologies Corporation | Geared turbofan gas turbine engine architecture |
US10287914B2 (en) | 2012-01-31 | 2019-05-14 | United Technologies Corporation | Gas turbine engine with high speed low pressure turbine section and bearing support features |
US20150345426A1 (en) | 2012-01-31 | 2015-12-03 | United Technologies Corporation | Geared turbofan gas turbine engine architecture |
US10125693B2 (en) | 2012-04-02 | 2018-11-13 | United Technologies Corporation | Geared turbofan engine with power density range |
CA2923331A1 (en) * | 2015-03-19 | 2016-09-19 | David P. Houston | Geared turbofan gas turbine engine architecture |
EP3633145A1 (en) * | 2018-10-04 | 2020-04-08 | Rolls-Royce plc | Reduced stress in compressor disc |
US11608750B2 (en) * | 2021-01-12 | 2023-03-21 | Raytheon Technologies Corporation | Airfoil attachment for turbine rotor |
Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5292385A (en) * | 1991-12-18 | 1994-03-08 | Alliedsignal Inc. | Turbine rotor having improved rim durability |
EP1424465A1 (en) * | 2001-09-03 | 2004-06-02 | Mitsubishi Heavy Industries, Ltd. | Hybrid rotor, method of manufacturing the hybrid rotor, and gas turbine |
US7241111B2 (en) * | 2003-07-28 | 2007-07-10 | United Technologies Corporation | Contoured disk bore |
EP1927722A1 (en) * | 2006-11-28 | 2008-06-04 | General Electric Company | Rotary assembly components and methods of fabricating such components |
US20090208339A1 (en) * | 2008-02-15 | 2009-08-20 | United Technologies Corporation | Blade root stress relief |
Family Cites Families (20)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2080425A (en) * | 1933-02-10 | 1937-05-18 | Milo Ab | Turbine |
JPS61234207A (en) * | 1985-04-10 | 1986-10-18 | Toshiba Corp | Method for extracting turbine blade |
EP0237170B1 (en) * | 1986-02-05 | 1994-05-11 | Hitachi, Ltd. | Heat resistant steel and gas turbine composed of the same |
FR2607866B1 (en) * | 1986-12-03 | 1991-04-12 | Snecma | FIXING AXES OF TURBOMACHINE ROTORS, MOUNTING METHOD AND ROTORS THUS MOUNTED |
US4784572A (en) | 1987-10-14 | 1988-11-15 | United Technologies Corporation | Circumferentially bonded rotor |
US4836750A (en) | 1988-06-15 | 1989-06-06 | Pratt & Whitney Canada Inc. | Rotor assembly |
US5067876A (en) * | 1990-03-29 | 1991-11-26 | General Electric Company | Gas turbine bladed disk |
US5632600A (en) | 1995-12-22 | 1997-05-27 | General Electric Company | Reinforced rotor disk assembly |
JP3149774B2 (en) | 1996-03-19 | 2001-03-26 | 株式会社日立製作所 | Gas turbine rotor |
US6019580A (en) | 1998-02-23 | 2000-02-01 | Alliedsignal Inc. | Turbine blade attachment stress reduction rings |
US6183641B1 (en) * | 1999-01-08 | 2001-02-06 | Fantom Technologies Inc. | Prandtl layer turbine |
EP1033476B1 (en) | 1999-03-03 | 2006-09-13 | General Electric Company | Heat exchange flow circuit for a turbine rotor |
GB2387203B (en) | 2002-04-02 | 2005-10-05 | Rolls Royce Plc | Rotor disc for gas turbine engine |
US6749400B2 (en) | 2002-08-29 | 2004-06-15 | General Electric Company | Gas turbine engine disk rim with axially cutback and circumferentially skewed cooling air slots |
EP1614857A1 (en) | 2004-07-05 | 2006-01-11 | Siemens Aktiengesellschaft | Turbomachine with a rotor comprising at least one drilled disc |
DE102004042295A1 (en) * | 2004-09-01 | 2006-03-02 | Mtu Aero Engines Gmbh | Rotor for an engine |
US8517666B2 (en) * | 2005-09-12 | 2013-08-27 | United Technologies Corporation | Turbine cooling air sealing |
US7578656B2 (en) | 2005-12-20 | 2009-08-25 | General Electric Company | High pressure turbine disk hub with reduced axial stress and method |
US8844265B2 (en) * | 2007-08-01 | 2014-09-30 | United Technologies Corporation | Turbine section of high bypass turbofan |
US8074440B2 (en) * | 2007-08-23 | 2011-12-13 | United Technologies Corporation | Gas turbine engine with axial movable fan variable area nozzle |
-
2012
- 2012-12-13 US US13/713,257 patent/US10119400B2/en not_active Ceased
-
2013
- 2013-09-24 WO PCT/US2013/061319 patent/WO2014092834A2/en active Application Filing
-
2020
- 2020-09-03 US US17/011,472 patent/USRE49382E1/en active Active
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5292385A (en) * | 1991-12-18 | 1994-03-08 | Alliedsignal Inc. | Turbine rotor having improved rim durability |
EP1424465A1 (en) * | 2001-09-03 | 2004-06-02 | Mitsubishi Heavy Industries, Ltd. | Hybrid rotor, method of manufacturing the hybrid rotor, and gas turbine |
US7241111B2 (en) * | 2003-07-28 | 2007-07-10 | United Technologies Corporation | Contoured disk bore |
EP1927722A1 (en) * | 2006-11-28 | 2008-06-04 | General Electric Company | Rotary assembly components and methods of fabricating such components |
US20090208339A1 (en) * | 2008-02-15 | 2009-08-20 | United Technologies Corporation | Blade root stress relief |
Also Published As
Publication number | Publication date |
---|---|
USRE49382E1 (en) | 2023-01-24 |
US10119400B2 (en) | 2018-11-06 |
WO2014092834A3 (en) | 2014-08-28 |
US20140109548A1 (en) | 2014-04-24 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
USRE49382E1 (en) | High pressure rotor disk | |
CA2933432C (en) | Geared turbofan gas turbine engine architecture | |
US10808543B2 (en) | Rotors with modulus mistuned airfoils | |
EP2971612B1 (en) | Engine mid-turbine frame transfer tube for low pressure turbine case cooling | |
EP3312080A1 (en) | Geared turbofan gas turbine engine architecture | |
EP3312404A1 (en) | Geared turbofan gas turbine engine architecture | |
EP3269965B1 (en) | Geared gas turbine engine | |
WO2013116257A1 (en) | Geared turbofan gas turbine engine architecture | |
EP2809940A1 (en) | Geared turbofan gas turbine engine architecture | |
EP3064711A1 (en) | Component, corresponding gas turbine engine and method | |
WO2014099634A2 (en) | Lightweight shrouded fan blade | |
US20130319011A1 (en) | Geared architecture carrier torque frame assembly | |
EP2935804B1 (en) | Gas turbine engine inner case including non-symmetrical bleed slots | |
EP2855874B1 (en) | Gas turbine engine with a counter rotating fan | |
EP3112613A1 (en) | Geared turbofan fan turbine engine architecture | |
EP2935791A1 (en) | Lightweight shrouded fan | |
EP3333365B1 (en) | Stator with support structure feature for tuned airfoil | |
EP3044424A2 (en) | Plug seal for gas turbine engine | |
WO2014092909A1 (en) | Multi-piece blade for gas turbine engine | |
EP3404215B1 (en) | Gas turbine engine with seal anti-rotation lock |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
122 | Ep: pct application non-entry in european phase |
Ref document number: 13861966 Country of ref document: EP Kind code of ref document: A2 |